Gas turbine engine ejector

ABSTRACT

An ejector comprises a primary nozzle having an annular wall forming part of an outer boundary of an exhaust portion of a primary flow path of a gas turbine engine. The annular wall has a downstream end defining a plurality of circumferentially distributed lobes. The ejector further comprises a secondary nozzle having an annular wall disposed about the primary nozzle, the primary nozzle and the secondary nozzle defining a secondary flow passage therebetween for channeling a secondary flow. The secondary nozzle defines a mixing zone downstream of an exit of the primary nozzle. A flow guide ring is mounted to the primary nozzle lobes. The ring has an aerodynamic surface extending from a leading edge to a trailing edge respectively disposed upstream and downstream of the exit of the primary nozzle. The aerodynamic surface of the ring is oriented to guide the high velocity primary flow into the mixing zone.

RELATED APPLICATION

This application is a divisional of U.S. patent application Ser. No.14/514,770, filed Oct. 15, 2014, which is a continuation-in-part of U.S.patent application Ser. No. 14/446,756 filed on Jul. 30, 2014, nowpatented under U.S. Pat. No. 9,745,919, issued Aug. 29, 2017 the contentof which is incorporated herein by reference.

TECHNICAL FIELD

The application relates generally to aircraft gas turbine engines and,more particularly, to aft section of the engine including an ejector.

BACKGROUND OF THE ART

In gas turbine engines, hot high velocity air exits from the turbinethrough the core gas path. The exhaust gases may be constrained by anexhaust case section in the form of a corrugated annular case extensionhaving lobes. Turbofan engines generally use exhaust mixers in order toincrease the mixing of the high and low velocity exhaust gas flows.Turbo-shaft and turbo-prop engines may be provided with similar devicessometimes referred to as ejectors. Exhaust mixers/ejectors mayexperience thermal variation and/or radial deflection due to exposure tothe high and low velocity flows. In addition, exhaust ejector/mixers maybe prone to vibrations, which have negative consequences for thesurrounding hardware. As such, it is generally desirable to increase thestiffness or rigidity of the exhaust case. Various configurations ofexhaust ejector/mixers have been proposed to date in order to try toincrease the stiffness or reduce deflection thereof.

Also, the aerodynamic performance of ejectors is often limited by theability of the primary flow to entrain the secondary cooling flow.Increasing the ejector capacity of pumping secondary mass flow wouldalso be desirable from an aerodynamic point of view.

SUMMARY

In one aspect, there is provided an ejector for a gas turbine engine ofthe type having a main axis and a primary flow passage channeling a highvelocity primary flow, the ejector comprising: a primary nozzle havingan annular wall forming part of an outer boundary of an exhaust portionof the primary flow passage, the annular wall having a downstream enddefining a plurality of circumferentially distributed radially innerlobes; a secondary nozzle having an annular wall disposed about theprimary nozzle, the primary nozzle and the secondary nozzle defining asecondary flow passage therebetween for channeling a secondary flow, thesecondary nozzle defining a mixing zone downstream of an exit of theprimary nozzle where the high velocity primary flow and the secondaryflow mix together; and a flow guide ring mounted to thecircumferentially distributed lobes in the exhaust portion of theprimary flow passage centrally about the main axis of the engine, theflow guide ring having an aerodynamic surface extending from a leadingedge to a trailing edge respectively disposed upstream and downstream ofthe exit of the primary nozzle, the aerodynamic surface being orientedto guide the high velocity primary flow into the mixing zone.

In accordance with another general aspect, there is provided a gasturbine engine having an engine casing enclosing a compressor section, acombustor and a turbine section defining a main gas path seriallyextending therethrough along a main axis of the engine, and comprising:an ejector projecting from an aft end of the engine casing axiallydownstream from an engine center body forming an aft end portion of aninner boundary of the main gas path, the ejector comprising a primarynozzle having an annular wall forming an outer boundary of the main gaspath for guiding a primary flow, the annular wall having a downstreamend defining a plurality of circumferentially distributed lobes, and aflow guide ring mounted to the circumferentially distributed lobes inthe main gas path centrally about the main axis and downstream of theengine center body; the flow guide ring having an aerodynamic surfaceconfigured to minimize diffusion of the primary flow towards the mainaxis of the engine.

BRIEF DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1 is a schematic cross-sectional view of a turbo-shaft gas turbineengine;

FIG. 2 is a rear isometric view of an exhaust ejector/mixer, having asupport member connected to the ejector/mixer lobes thereof, inaccordance with one embodiment of the present disclosure;

FIG. 3 is an enlarged fragmentary, isometric view of a lobe and supportmember according to FIG. 2;

FIG. 4 is a fragmentary rear isometric view an ejector/mixer, having asupport member connected to the lobes thereof, in accordance withanother embodiment;

FIG. 5 is an enlarged fragmentary, isometric view of a lobe and supportmember according to FIG. 4;

FIG. 6 is a schematic, axial cross section of a portion of theejector/mixer showing the main gas path, and the support member locatedand oriented in the gas path;

FIG. 7 is a schematic, radial cross section of a portion of theejector/mixer showing the hot main gas path and the induced cool air inthe lobes; and illustrating the relative location of the support member;

FIG. 8 is a rear isometric view of an ejector suited for mounting at theexhaust end of a turbo-prop or a turbo-shaft;

FIG. 9 is a schematic, axial cross section of a portion of the ejectorshown in FIG. 8; and

FIG. 10 is an enlarged radial cross section of a portion of the ejectorshown in FIG. 8 and illustrating the ejector lobe draft angle.

DETAILED DESCRIPTION

FIG. 1 illustrates a turbo-shaft gas turbine engine 10 of a typepreferably provided for use in subsonic flight, generally comprising inserial flow communication a compressor section 14 for pressurizing theair, a combustor 16 in which the compressed air is mixed with fuel andignited for generating an annular stream of hot combustion gases, and aturbine section 18 for extracting energy from the combustion gases. Thegas turbine engine 10 includes a core engine casing 20 which enclosesthe turbo machinery of the engine. The main air flow passes through thecore of the engine via a main gas path 26, which is circumscribed by thecore engine casing 20 and allows the flow to circulate through themultistage compressor 14, combustor 16 and turbine section 18 asdescribed above.

At the aft end of the engine 10, an engine center body 22 is centeredabout a longitudinal axis X of the engine 10, the engine center body 22being connected to an aft end of the turbine section 18. The enginecenter body can take the form of an exhaust cone depending on theapplication. The engine center body 22 has an outer surface, whichdefines an inner wall of the main gas path 26 so that the combustiongases flow therearound. An ejector/mixer 30 forms the outer wall of theaft end of the main gas path 26. As best seen in FIG. 2, theejector/mixer 30 includes a primary nozzle having an annular wall 34with a radial fastening ring or flange 32 upstream thereof. Thefastening ring 32 is adapted to be mechanically fastened to the aftportion 20 a (FIG. 1) of the casing 20.

Referring to FIGS. 2 and 3, in further detail, the annular wall 34 ofthe primary nozzle, includes and defines a plurality ofcircumferentially distributed radially inner lobes 36 extending axiallyrearwardly from a front frusto-conical portion of the annular wall 34 toa downstream edge 37, i.e. a trailing edge thereof. The lobes 36 includeside, radially-extending, walls 38 with a bight forming an arcuatetrough 40. The trough 40 presents a convex surface 41 on the radiallyinner or central side of the annular wall 34.

An annular support member includes an annular blade 42 extendingconcentrically about the longitudinal axis X of the engine 10. In theembodiment shown, the blade 42 comprises an annular longitudinal, flatbar. The blade 42 is interrupted only at form-fitting joint areas 44.The joint areas 44 are located on the blade 42 to correspond with theconvex surfaces 41 of the lobes 36. The joint areas 44 are curved sothat it complements the convex surface 41, as shown in FIG. 3. Thecurved joint area 44 is concave relative to the convex surface 41 of thelobe 36. The blade 42 is spaced radially outwardly and independent fromthe engine center body 22 and floats with respect thereto. The blade 42in one embodiment is a thin sheet metal strip capable of being welded tothe sheet metal forming the annular wall 34. In the embodiment shown inFIGS. 2 and 3, the thin sheet metal strip is formed into a continuousring.

As mentioned, the ejector/mixer 30 is solely connected to the engine 10at the aft end 20 a of the core engine casing 20, and so, theejector/mixer 30 is effectively cantilevered from the core engine casing20. This cantilevered configuration allows the lobes 36 of the exhaustejector/mixer 30 to vibrate at one or more modes in the engine operatingfrequency range, while remaining relatively stiff. In addition, thethermal variations in the exhaust mixer 30 due to the high and lowvelocity flows through the main gas path 26 may cause axial and radialdisplacements in the ejector/mixer 30, which can accordingly be absorbedby the exhaust ejector/mixer 30. Moreover, the downstream end 37 of theejector/mixer 30, which would otherwise be prone to deflection, isreinforced by the blade 42 which serves to increase the rigidity of theexhaust ejector/mixer 30 and thus inhibit movement at the downstream end37 thereof. By joining all the lobes 36 together with the blade 42, anymovement of the ejector/mixer 30 is reduced, as are the vibrationsthereof. In addition, by providing a blade 42 which is independent ofthe exhaust engine center body 22, i.e. it is free to move relativethereto such as to absorb any vibrations or thermal growth mismatchestherebetween. The blade 42 is able to accommodate any axial or radialdisplacements due to such thermal variations. As such, the ejector/mixer30 provides enhanced rigidity and may accommodate thermal variations,vibrations and other displacements, as required.

Another embodiment is shown in FIGS. 4 and 5. In this case, the blade ismade up of blade segments 142 a, 142 b . . . 142 n. Each segment has alength corresponding to the distance between the center lines of twoadjacent lobes 36. Each end of the segment terminates in a partiallyformed concave curve to complement part of the convex surface 41 of thelobe 36. For instance, as shown in FIG. 5, corresponding ends ofsegments 142 a and 142 b make-up the form fitting joint area 144.

The blade 42, 142 may be located at different axial positions along theconvex surfaces 41 of the lobe 36. FIG. 3 illustrates a blade 42 beingspaced upstream from the trailing edge 37, of the lobe 36. As shown inFIG. 5, the blade 142 is located at or slightly downstream from thetrailing edge 37, of the lobe 36. The blade 42, 142 may be fixed to theconvex surfaces 41 of the lobes 36 at joint areas 44, 144 using acombination of resistance, fusion or ball tack welding with a brazingapplication, or other types of suitable bonding that are well known inthe art.

The injector/mixer 30, in the present embodiment, acts to induce coolair, exterior of the engine casing 20, to be drawn radially inwardlythrough the lobes 36 to cool the mechanical parts of the injector/mixer30. As previously mentioned, the support member is often, according tothe prior art, subject to thermal stresses caused by the entrained coolair and of the hot air exiting the turbine 18. FIGS. 6 and 7 show thegases flow in the ejector/mixer 30. The blade 42, 142 is disposeddirectly in the main gas path 26 and is shaped to be laminar with theflow of the hot gases, as can be seen in both FIGS. 6 and 7. The blade42 is essentially exposed only to the hot gases flowing in the main gaspath 26. This reduces the thermal gradient in the blade 42, 142.

The embodiments described show a turbo-shaft engine. However, in thecase of a turbofan engine, cool air from the fan is directed to theejector/mixer 30 which in such a case would have inner and outeralternating lobes to best mix the hot gases with the cool air. U.S. Pat.No. 5,265,807 Steckbeck et al 1993; U.S. Pat. No. 7,677,026 Conete et al2010; and U.S. Pat. No. 8,739,513 Lefebvre et al 2014 describe exhaustmixers which are herewith incorporated by reference.

The above described embodiments provide an improved exhaustejector/mixer for a gas turbine engine where the thermal stresses on thesupport member are reduced for improved longevity.

It is noted that the ejector/mixer and the support member could be madeby additive manufacturing processes, such as direct metal lasersintering. Therefore, the ejector/mixer and the support member could bemade monolithically.

For some gas turbine engine applications, such as turbo shaft and turboprop applications, where the engine center body 22 ends axially upstreamof the turbine exhaust nozzle exit (see FIG. 9), the exhaust section isreferred to as an ejector. As will be seen hereinafter, for suchapplications, the support member may also act as a flow guide ring toguide the primary flow when leaving the primary nozzle and, thus,enhance the ejector aerodynamic performance.

FIG. 8 illustrates an ejector 200 comprising a primary nozzle 201, asecondary nozzle 203 concentrically mounted about the primary nozzle 201and a flow guide ring 205 concentrically mounted inside the primarynozzle 201.

As mentioned hereinbefore with respect to the embodiments shown in FIGS.1 to 7, the primary nozzle 201 has an annular wall 234 forming part ofthe outer boundary of an exhaust portion of the main or primary flowpassage of the engine. The annular wall 234 has a downstream end formedwith circumferentially distributed radially inner lobes 236. The flowguide ring 205 is attached to the radially inner surface of the lobevalleys as described herein above.

The secondary nozzle 203 has an annular bell-shaped wall extending fromthe engine compartment wall case (not shown) about the primary nozzle201. As best shown in FIG. 9, the primary nozzle 201 and the secondarynozzle 203 define a secondary flow path 207 therebetween for guiding asecondary flow of cooling air. The secondary nozzle 203 extends axiallydownstream of the primary nozzle 201 and defines a mixing zone 209 atthe exit of the primary nozzle 201 where the high velocity primary flowmixes with the secondary flow.

Referring conjointly to FIGS. 8 and 9, it can be appreciated thatprimary nozzle 201 of the ejector 200 extends axially downstream of theengine center body 22 (i.e. the inner boundary of the primary flowpassage ends upstream of its outer boundary). As a result, the primaryflow tends to diffuse towards the engine centerline downstream of theend of the center body 22.

The addition of a properly designed flow guide ring 205 can prevent theannular high momentum primary flow from diffusing and guide the flowthrough the annular zone between the flow guide ring 205 and the primarynozzle 201 where the primary and secondary flows mix. Due to this fact,the capacity of pumping secondary mass flow may be improved.

According to the embodiment illustrated in FIGS. 8 and 9, the flow guidering 205 has a cone shape with a proper angle (P1) with respect toengine axis (see FIG. 9). This is to ensure that the primary flow iswell guided without separation when leaving the primary nozzle andentering the mixing zone. The ring cone draft angle (P1) may be in therange of about 0° to about 10° and is preferably about 5°. Depending onthe application, the flow guide ring 205 could be cylindrical or airfoilas well.

As shown in FIG. 9, the flow guide ring 205 has an aerodynamic surfaceextending axially from a leading edge 205 a to a trailing edge 205 b.According to the illustrated embodiment, the leading edge 205 a and thetrailing edge 205 b are respectively disposed upstream and downstream ofthe primary nozzle exit to properly guide the primary flow leaving theprimary nozzle 201 into the mixing zone 209. According to theillustrated embodiment, the flow guide ring 205 projects out of theprimary nozzle 201 or extends downstream from the exit of the primarynozzle 201 by a distance (P5) for extended flow guidance in the mixingzone and avoidance of flow separation across the ring 205. For aparticular application, the distance (P5) is in the range of about 0 toabout 1 inch.

The length (P2) of the guide ring 205 and its axial installationposition (P3) relative to the end of the center body 22 may alsoinfluence the aerodynamic performance of the ejector 200. It isunderstood that (P2) and (P3) can be optimized depending on differentapplications. According to a particular application, the ring length(P2) is in the range of about 0.5 to about 2 inches and the ring 205 isinstalled axially at the primary nozzle exit.

The radial installation position of the guide ring (P4) may varydepending on various conditions. According to the illustratedembodiment, the ring 205 is installed at the lobe valley. It is alsounderstood that the lobe design and the number of lobes 236 may varydepending on the applications. According to the illustrated embodiment,the lobes 236 have a draft angle (P6) of about 0° to about 5° (FIG. 10).Such a small draft angle can help prevent reverse back secondary flow.The number of lobes may vary depending on the size of the engine. Forthe exemplified application given above, the number of lobes may rangebetween 8 and 10.

Various permutations of the above parameters of the flow guide ring canbe used to improve the ejector pumping capacity.

The above description is meant to be exemplary only, and one skilled inthe art will recognize that changes may be made to the embodimentsdescribed without departing from the scope of the invention disclosed.For example, the invention may be used with various types of gas turbineengines where cool and hot gases may simultaneously be in contact withthe machinery involved. Still other modifications which fall within thescope of the present invention will be apparent to those skilled in theart, in light of a review of this disclosure, and such modifications areintended to fall within the appended claims.

1. A gas turbine engine having an engine casing enclosing a compressorsection, a combustor and a turbine section defining a main gas pathserially extending therethrough along a main axis of the engine, andcomprising: an ejector projecting from an aft end of the engine casingaxially downstream from an engine center body forming an aft end portionof an inner boundary of the main gas path, the ejector comprising aprimary nozzle having an annular wall forming an outer boundary of themain gas path for guiding a primary flow, the annular wall having adownstream end defining a plurality of circumferentially distributedlobes, and a flow guide ring mounted to the circumferentiallydistributed lobes in the main gas path centrally about the main axis anddownstream of the engine center body; the flow guide ring having anaerodynamic surface configured to minimize diffusion of the primary flowtowards the main axis of the engine.
 2. The gas turbine engine accordingto claim 1, wherein the ejector further comprises a secondary nozzleconcentrically mounted about the primary nozzle, the primary nozzle andthe secondary nozzle defining a secondary flow passage therebetween forchanneling a secondary flow, the secondary nozzle circumscribing amixing zone downstream of an exit of the primary nozzle, and wherein theflow guide ring is configured and disposed to guide the primary flow atits exit from the primary nozzle into the mixing zone.
 3. The gasturbine engine according to claim 1, wherein the aerodynamic surfaceconverges radially inwardly towards the main axis of the engine in adownstream direction.
 4. The gas turbine engine according to claim 1,wherein the flow guide ring is a circumferentially continuous one-piecemetallic strip mounted to a radially innermost surface of thecircumferentially distributed lobes.
 5. The gas turbine engine accordingto claim 1, wherein the engine is a turbo-shaft engine or a turbo-propengine.
 6. The gas turbine engine according to claim 4, wherein the flowguide ring has a conical shape.
 7. The gas turbine engine according toclaim 1, wherein the flow guide ring extends axially downstream from theexit of the primary nozzle by a distance P5, and wherein P5 is equal toor less than about 1 inch.
 8. The gas turbine engine according to claim1, wherein the aerodynamic surface of the flow guide ring extends at anangle to the main axis, the angle being equal to or less than 10degrees.
 9. The gas turbine engine according to claim 8, wherein theangle is about 5 degrees.
 10. The gas turbine engine according to claim1, wherein the lobes have a draft angle comprised between about 0° andabout 5°.
 11. The gas turbine engine according to claim 1, wherein theflow guide ring has an axial length P2 comprised between about 0.5inches and about 2 inches.